Integrated heat management for hybrid propulsion

This application claims priority from U.S. Provisional Patent Application No. 62/575,599 filed Oct. 23, 2017, the entire contents of which are incorporated by reference herein.

The application relates generally to hybrid electric aircrafts and, more particularly, to a heat management system for power plants of such aircrafts.

Heretofore, the cooling of hybrid electric aircraft propulsion power plants has proven to be challenging. Typically, individual cooling systems are provided for each heat generating component of the power plant, thereby resulting in extra weight and integration complexity.

There is thus a need for a new heat management system suited for hybrid electric aircrafts.

In accordance with a general aspect, there is provided a power plant for a hybrid electric aircraft, the power plant comprising: at least one electric propulsor having a nacelle housing a fan for generating thrust for the aircraft; a generator for supplying power to the at least one electric propulsor; at least one combustion engine operatively connected to the generator; a heat exchanger connected in heat exchange relationship with both the generator and the at least one combustion engine, the heat exchanger being provided at an outer duct wall of the nacelle of the at least one electric propulsor.

In accordance with another general aspect there is provided an aircraft comprising: an electric propulsor having a nacelle housing a fan driven by an electric motor; a source of power for supplying power to the electric propulsor, the source of power including a battery pack and a generator; a combustion engine operatively connected to the generator; and a common cooling system for the battery pack, the generator and the combustion engine, the common cooling system comprising a heat exchanger integrated to the nacelle of the electric propulsor downstream of the fan.

In accordance with a still general aspect, there is provided a method of managing heat generated by a power plant of an aircraft having electric propulsors powered at least in part by a generator operatively connected to a combustion engine, the method comprising: withdrawing heat from the generator and the combustion engine and dissipating the heat in ambient air outside the aircraft via a heat exchanger operatively connected to both the generator and the combustion engine.

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic isometric view of a hybrid electric aircraft having a pair of electric propulsors mounted on opposed sides of the aircraft fuselage;

FIG. 2 is a schematic view of a power plant for supplying power to the electric propulsors of the aircraft; and

FIG. 3 is a schematic cross-section view of one of the electric propulsors of the aircraft and illustrating an example of an integration of a heat management system into the nacelle of the propulsor.

Referring to FIG. 1 , there is shown an example of a hybrid electric aircraft 10 including a pair of electric propulsors 14 mounted on opposed sides of the aircraft fuselage 11 for generating thrust for the aircraft 10 . As show in FIG. 3 , each electric propulsor 14 may be provided in the form of a ducted fan including a nacelle 20 housing a fan 22 driven in rotation by an electric motor 13 mounted inside a centerbody 21 downstream of the fan 22 . An annular air passage 24 is defined between a radially inner surface of the nacelle 20 and a radially outer surface of the centerboby 21 for ducting the external air drawn by the fan 22 . As shown in FIG. 3 , the air passage 24 may have a converging cross-section profile or nozzle shape to further accelerate the stream of air drawn by the fan 22 in order to generate greater thrust for the aircraft 10 .

As shown in FIG. 2 , various sources of power may be used for supplying electric energy to the electric propulsors 14 . For instance, the source of power may comprises a battery pack 12 or another suitable energy storage unit and/or an electrical generator 16 .

The electrical generation can be provided by a combustion engine driving the generator 16 . The combustion engine can adopt various forms, such as a conventional gas turbine engine, an internal combustion engine (ICE) having a variable volume combustion chamber (e.g. a piston engine or a Wankel engine), or a compounded engine (e.g. an ICE integrated to a gas turbine engine). According to one embodiment, the electrical generation is provided at least in part by a conventional gas turbine engine 17 driving generator 16 , which, in turn, supply electric energy to the electric motors 13 of the propulsors 14 . According to another embodiment, the generator 16 could be solely driven by an internal combustion engine (ICE) 19 , such as a piston engine or a Wankel engine, having a variable volume combustion chamber. According to a further variant, a compounded engine comprising a gas turbine engine and an ICE may be provided to drive the generator 16 . It is also understood that the generator 16 could be operatively connected to both an ICE and a gas turbine engine. Various combination of fuel powered engines are contemplated.

In operation, the various elements of the power plant, including the batteries 12 , the generator 16 , the gas turbine engine 17 , the ICE 19 and/or the compounded engine need to be cooled. While conventional gas turbine engines are typically air cooled using a small % amount of the main engine air, the batteries 12 , the generator 16 and the ICEs are typically cooled by separate liquid-based cooling systems. Such systems typically comprise a heat exchanger defining a coolant circuit through which a coolant (e.g. a liquid or gas) is circulated to remove heat from the parts to be cooled.

It is herein proposed to combine all the cooling systems into one and to integrate the combined cooling system into the aircraft external envelope, which is exposed to ambient air. For instance, in the case of the exemplified fixed wing aircraft equipped with electric propulsors 14 , heat exchangers 18 could be integrated to the propulsors nacelle 20 in order to effectively dissipate the heat picked up by the liquid coolant from batteries 12 , the generator 16 and the ICE 19 . More particularly, in the case of a ducted fan having a fan 22 mounted in a nacelle 20 including an outer duct wall circumscribing an air passage 24 as shown in FIG. 3 , a heat exchanger 18 could be integrated to the outer duct wall of the nacelle 20 downstream of the fan 22 to dissipate the combined heat of the ICE, batteries 12 and/or electrical generator 16 into the stream of air flowing through the air passage 24 , thereby increasing the energy of the air propelled by the fan 22 and, thus, improving the performance of the propulsors 14 . This can, for instance, be achieved by integrating an annular liquid-air heat exchanger in the outer duct wall of the air passage 24 . In operation, the liquid coolant is circulated to pick up heat from the batteries 12 , the generator 16 and the ICE 19 and is then routed through the heat exchanger 18 where heat from the liquid coolant is transferred to ambient air flowing over the nacelle and to the air flowing through the air passage 24 of the ducted propulsors 14 for generating thrust for the aircraft 10 .

As schematically shown in FIG. 3 , an additional heat exchanger 26 could be integrated in a front lip portion of the propulsor nacelle 20 upstream of the fan 22 to perform a lip de-icing function. Indeed, a portion of the heat withdrawn from the batteries 12 , the generator 16 and/or the ICE by the liquid coolant could be re-used to de-ice the nacelle inlet lip 26 without any impact on engine specific fuel consumption (SFC). This would be an improvement over traditional de-icing system which uses either engine bleed or power extraction with a SFC increase.

Also, the exhaust gases of the ICE 19 , gas turbine engine 17 and/or compounded engine of the aircraft power plant could be discharged in the air passage 24 of the propulsor nacelle 20 downstream of the heat exchanger 18 to further energize the air discharged from the electric propulsors 14 (i.e. the airflow stream of the fan according to the illustrated example). Effects A, B & C shown in FIG. 3 could be used in combination or individually as a way to optimize the integration of a liquid cooled engine into hybrid electric aircrafts.

The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For instance, only heat withdrawn from the liquid cooled engine could be dissipated through the propulsor nacelle. Alternatively, when the generator is driven by a conventional gas turbine engine, only the heat generated by the generator and/or the batteries could be dissipated via heat exchangers integrated to the nacelle of the propulsors. Also it is understood that it is not necessarily all the heat generated by the batteries the generator and the ICE that is dissipated through the propulsor nacelles. Indeed, additional heat exchangers could be provided on aircraft fuselage if the amount of heat to dissipate is too important. Lastly, it is understood that the aircraft could comprises any number of electric propulsors. Any modifications which fall within the scope of the present disclosure will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.